The present invention relates generally to airfoil structures for turbine engines, and more particularly to airfoil structures including a composite core.
Composite blades developed for commercial aircraft engine fan blades may be constructed of laminated carbon/epoxy “prepreg” material. A “prepreg” is a layer of fibers, for example carbon fibers impregnated with a resin and arranged to form a lamina or a cloth. Prepreg layers may be layered and cured to form a composite structure or laminate. The laminates may experience interlaminar separation under certain circumstances. Moreover, when laminated fan blades are subject to high energy impacts (e.g., birds, or other foreign objects), the interlaminar separation can result in delamination and a reduction in the blade's structural integrity.
The shear stresses that may tend to delaminate the blade structure are generated when the composite blade is subjected to high twisting and bending loads. These loads normally result from impacts which often occur on the leading edge of the blade. When the blade is subjected to an impact, the peak shear stresses tend to be transmitted to the middle of the blade, as well as the leading and trailing edges.
Previous attempts to improve resistance to delamination of composite fan blades have involved, for example, stitching a full-sized “all prepreg” blade before cure, or by using 3D woven structures. 3-D type woven structures have been researched extensively to increase the delamination resistance and decrease the damage area during the impact, where a certain number of reinforcement fiber tows were woven in a through-thickness direction or partially through-thickness direction. However 3-D woven based blades may have lower stiffness and initial failure strain.
Thus, it would be desirable to provide improved structures for composite fan blades that provide high impact resistance, superior damage tolerance and less complex manufacturing.